This invention relates generally to low thrust rocket propulsion thrusters and more particularly to bipropellant chemical thrusters for in-space satellite attitude and orbit control and in-space vehicle propulsion.
In-space propulsion thrusters are used to maneuver spacecraft (e.g. a satellite or space vehicle) after a launch vehicle has delivered it to the upper atmosphere. In general, the primary objective of a space propulsion thruster is to place the spacecraft into its intended orbit or maintain the spacecraft's proper position while in orbit. Specifically, onboard thrusters are used for orbit transfer; attitude pointing and control so that a spacecraft is correctly pointing towards the Earth, Sun or an astronomical object of interest; orbit altitude control; and station keeping. Thrust attitude control allows spacecraft to control the angular position while in orbit, which may be required for various sensors, transponders or other spacecraft hardware. Thruster systems must be able to operate in various propulsion modes, including short engine pulses to long duration steady-state firings depending on the mission requirements.
While in space, the purpose of the propulsion thruster is to change the velocity of the spacecraft. Because this is more difficult for larger spacecraft, propulsion thruster designs normally work with momentum (mv). The rate of change in momentum is referred to as a force (F=d/dt(mv)). Furthermore, acceleration is the rate of change in velocity (a=d/dt(v)). The goal of in-space thrusters is to create a force over a period of time, which is called an impulse (FΔt=mv). A spacecraft can be propelled to a specific velocity by applying a small acceleration over a long period of time (Δt), or alternatively, a large acceleration over a short period of time (v=aΔt). Similarly, a given impulse can be achieved with a large force over a short period of time or conversely with a small force over a longer time. This means that for maneuvering in space, a propulsion system that produces a very small acceleration but over a longer time can generate the same impulse as a propulsion system that produces a large acceleration, but over a short period of time (FΔt=maΔt).
When evaluating the efficiency of a propulsion system, designers normally refer to the force or velocity produced relative to the amount of mass that has to be carried along with the rocket or thruster that is irretrievably consumed when used to generate thrust (i.e., finite amount of propellant available for a given mission). The performance of a rocket engine is typically characterized by the specific impulse, Isp, which is the ratio of the engine thrust, F, to the mass flow rate ejected, {dot over (m)}. Thus, Isp=F/({dot over (m)})=ve/gc, with ve defined as the exhaust velocity, and gc the earth's gravitational acceleration, with units of seconds. When the thrust and the mass flow rate remain constant throughout the burning of the propellant, the specific impulse is the time during which the rocket engine provides a thrust equal to the amount of propellant consumed. Thus, to maximize thrust for a given amount of propellant consumption that is carried onboard the spacecraft requires a high specific impulse. For a given rocket engine, the specific impulse has a different value on the ground versus in the vacuum of space, due to the absence of atmospheric pressure. Hence, it is important to differentiate between specific impulse at sea level or in a vacuum.
Chemical propulsion thruster systems for spacecraft usually employ liquid reactants as the energy source. The propellant can be a single reactant (monopropellant) or a combination of liquid fuel and oxidizer (bipropellant). For a monopropellant system the most common propellant is hydrazine. Generally, for small thruster designs hydrazine is passed through a catalyst bed. As a result, thrust is produced by the decomposition of the propellant and catalyst into ammonia, nitrogen and hydrogen at a temperature of about 1300° F. Ignition of monopropellants can be produced thermally or by a catalytic material. Monopropellant propulsion systems are usually employed for attitude control and station-keeping since they are well suited to produce short duration pulses of thrust from less than a pound up to about 5 lbf with an accompanying Isp of about 230 seconds. Short duration pulses can range from about 0.01 or 0.02 seconds to about 0.10 seconds, and as a result the specific impulse can lose anywhere from about 50% to about 75% or 85% of the theoretical impulse value, respectively. Thus, monopropellant thrust systems typically have low Isp values. Since hydrazine is a highly toxic fuel (due to its vapors) and capable of exploding at 450° F., special safety features are required during use. When properly sealed, however, hydrazine stores well making it a widely used propellant.
For most bipropellant systems, nitrogen tetroxide is typically utilized as the oxidizer and either hydrazine or monomethyl hydrazine (MMH) is employed as the fuel. The reactants are hypergolic, meaning the fuel burns spontaneously upon contact with the oxidizer, hence facilitating ignition under vacuum conditions and in the pulsed mode of operation. Additionally, non-hypergolic bipropellants require some form of an ignition system to initiate combustion. Use of hypergolic propellants eliminates the need for an ignition system when multiple re-starts are required. The specific impulse of such a chemical propulsion thruster system would typically range from approximately 290 to 310 seconds with a thrust range typically between 90 lbf to about 140 lbf. Such characteristics make hypergolic propellants well-suited for final orbit apogee insertion after initial drop-off by the launch vehicle. A smaller version of this thruster design could also be used for attitude control. Again, hydrazine or MMH vapors are extremely toxic, requiring special handling and the use of two propellants somewhat complicates the propellant management for on-board spacecraft.
For a typical spacecraft operating in Earth orbit, the weight of the propulsion thrust system, including onboard propellants, can range from 10% to 20% of the total spacecraft weight, and up to 40% to 50% if the spacecraft is required to significantly alter its orbit. As a result, technology improvements have focused on achieving higher specific impulse Isp, since about 90% of the thrust propulsion system consists of propellants. Most recent improvements in rocket thruster technology have concentrated on increasing the allowable operating temperature of the combustion chamber to achieve small reproducible impulse without affecting the overall specific impulse. However, the general goal of chemical thruster technology is to develop high specific impulse rocket systems. For small thruster systems that may use hypergolic, advanced, or traditional rocket propellants, high specific impulse rocket systems are achieved by increasing combustion and propulsive efficiencies and increasing performance across a broad spectrum of thrust levels (less than about 5 lbf to about 250 lbf and upwards to about 500 lbf). Improvements in high-temperature materials for combustor/nozzle components also increase the specific impulse of a rocket thrust system. Thus, it is typically always desirable to increase the specific impulse (currently to above 350 seconds), minimize rocket weight and mass, operate radiation cooled rockets at arbitrary propellant mixture ratios with all onboard propellant options and reduce overall costs. To put the specific impulse goal in perspective, the SSME (Space Shuttle Main Engine) rocket engine using liquid hydrogen/liquid oxygen, has a very high vacuum specific impulse of 452 seconds and a vacuum thrust level of 491,000 lbf. This very high efficiency is achieved by utilizing a staged combustion cycle, whereby a portion of the propellants that are partially combusted, at a fuel-rich mixture ratio, is used to drive the high pressure turbo-pump prior to undergoing combustion in the main combustion chamber. This type of rocket engine is much too complicated and cannot be miniaturized for implementing into small spacecraft thruster systems.
Recently, aerovortical swirl-dump combustion (ASC) technology has been developed and introduced into airbreathing, ramjet, combined-cycle, and rocket propulsion applications to improve engine performance. The key feature of the swirl-dump combustion technology is the swirl generator. The swirl generator with a dump-combustor design is able to obtain near complete combustion of the liquid propellants over a wide range of mixture ratios and within very short combustor lengths and diameters. High propulsion performance has been test demonstrated in a combustor-convergent nozzle length to diameter ratio (L/d) of 1.6; while analysis shows that this L/d can be further reduced down to 1.0 or less with equally high engine performance. Furthermore, the swirl generator has no moving parts so the complexity of the engine and production cost is kept low. The swirl generator introduces a swirling flowfield through the use of a stationary vane design into which the liquid fuel and/or oxidizer propellants are introduced. Each swirl vane imparts tangential and radial velocities into the combustion constituents, thereby producing a highly turbulent three-dimensional flowfield in the combustor. The high turbulence scale and intensity in this swirling aeroflow structure rapidly and efficiently improves atomization, vaporization, mixing and burning of the injected fuel and oxidizer propellants. In addition, the swirl generator design improves flame propagation and spreading, operability range and combustion stability. All of these features result in a very high combustion efficiency and high performance in short combustor lengths over wide flammability limits. Thus, the size and weight of an ultra-compact rocket engine thruster can be significantly reduced, while maintaining high propulsion performance if it could be combined with swirl combustion technology.
However, current swirl vane designs are limited in their applicability to ultra-compact rocket engine designs due to their low-end size limitation. For ultra-compact rocket engine designs that are, for example, less than about two inches in diameter, it is very difficult to fabricate, integrate and assemble individual swirl vanes into the vane pack. This hampers practical application of swirl combustion technology into the ultra-compact rocket engine designs. Thus, there is a need for smaller, lighter and better performing ultra-compact rocket engine designs suitable for spacecraft applications. Specifically, there is a need for swirl combustion technology suitable for use in ultra-compact rocket engine designs.